Gas turbine engine cooling system control

ABSTRACT

A gas turbine engine includes a core cowl and a core contained within the core cowl. The core includes a compressor in fluid communication with a downstream combustor and a downstream turbine, the compressor including a compressor bleed port, wherein an undercowl space is defined between the core cowl and the core. The gas turbine engine further includes a cooling duct disposed at least partially in the undercowl space and having an inlet and an outlet, wherein the cooling duct is in fluid communication with a source of cooling air and is further in fluid communication with the compressor bleed port; a valve assembly including at least one valve disposed in the cooling duct; and a cooling blower disposed within the engine and operable to move an air flow from the inlet of the cooling duct towards the outlet of the cooling duct and through the compressor bleed port.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation application of U.S. application Ser.No. 17/160,742, filed Jan. 28, 2021, which is a non-provisionalapplication and is hereby incorporated by reference in its entirety.

FIELD

The present disclosure relates to a gas turbine engine having a coolingsystem, such as a reverse bleed system, and a method of operating thesame.

BACKGROUND

During normal operations, temperatures of gas turbine engine componentsare maintained within allowable limits by a plurality of coolingprocesses that transfer heat from the components to one or more heatsinks. When the engine is shutdown, most cooling systems no longeroperate. Residual heat in certain engine components (i.e., “soakback”)can increase the temperature of other engine components beyond allowablelimits, and further may unevenly heat other components sometimescreating a “bow” in the components, also referred to as a “rotor bow.”

A particular concern is the formation of carbon (or “coke”) deposits infuel carrying components including fuel nozzles when a hydrocarbon fuel(liquid or gas) is exposed to high temperatures in the presence ofoxygen. Some known methods of mitigating coking include rotating therotor after engine shutdown (i.e. “motoring”) or purging the engine withforced air provided from an auxiliary power unit (“APU”), ground powerunit (“GPU”), or air conditioning unit after shutdown.

Similarly, a concern with rotor bow is prematurely wearing out seals andclearances, e.g., within the compressor section of the gas turbineengine causing lower efficiency and more frequent repairs. Some knownmethods of fixing a bowed rotor include relatively slow motoring of theengine for an extended period of time prior to restarting the engine toredistribute the heat and/or cool the components.

One problem with these methods is that they require resources such aselectrical power, fuel, external equipment, and/or logistical supportthat may be unavailable or impractical, and further may increase astartup time for the gas turbine engine.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary aspect of the present disclosure, a method is providedfor operating a gas turbine engine. The method includes: determiningdata indicative of an operation of a cooling system of the gas turbineengine during a shutdown of the gas turbine engine, following theshutdown of the gas turbine engine, or both; and modifying a startupschedule of the gas turbine engine in response to the determined dataindicative of the operation of the cooling system of the gas turbineengine.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a cross-sectional, schematic view of a gas turbine engine inaccordance with an exemplary aspect of the present disclosure includinga reverse bleed system in accordance with an exemplary aspect of thepresent disclosure.

FIG. 2 is a schematic, close-up, sectional view of the exemplary reversebleed system suitable within the gas turbine engine of FIG. 1 .

FIG. 3 is a schematic perspective view of the gas turbine engine of FIG.1 mounted to an aircraft.

FIG. 4 is flow diagram of a method for operating a gas turbine engine inaccordance with the present disclosure.

FIG. 5 is a graph depicting certain parameters of an engine operated inaccordance with the method of FIG. 4 .

FIG. 6 is a control scheme in accordance with the present disclosure.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations. Additionally, unlessspecifically identified otherwise, all embodiments described hereinshould be considered exemplary.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 1, 2, 4,10, 15, or 20 percent margin. These approximating margins may apply to asingle value, either or both endpoints defining numerical ranges, and/orthe margin for ranges between endpoints.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

In one exemplary aspect of the present disclosure, a method foroperating a gas turbine engine is provided, whereby certain dataregarding operation of an air cooling system for a gas turbine engineare determined when the air cooling system is operated following ashutdown of the gas turbine engine. The air cooling system may be areverse bleed system, and may be configured to reduce or minimize a cokeformation within, e.g., fuel nozzles of a combustor of the engine, andmay further be configured to reduce or minimize a rotor bow conditionwithin the engine. The data determined may be based on sensed datastored in non-volatile memory of the engine controller prior to theengine controller shutting down. The data determined may indicatewhether or not the air cooling system was operating properly, wasprematurely shutdown, etc.

In certain aspects of the present disclosure, the method of the presentdisclosure may modify a startup sequence for a subsequent startupoperation of the gas turbine engine based on the data determinedregarding the operation of the air cooling system. For example, if thedata indicated that the air cooling system was operating properly, thenthe modification may be to shorten a motoring of the engine prior toaccelerating the engine to a light-off rotational speed, as such mayindicate that the engine was properly cooled and is not experiencingrotor bow. By contrast, if the data indicated that the air coolingsystem was not operating properly, then the modification may be tolengthen the motoring of the engine prior to accelerating the engine tothe light-off rotational speed, as such may indicate that the engine wasnot properly cooled and may be experiencing rotor bow.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 depicts an exemplary gasturbine engine 10 defining an axial direction A (and centerline axis 11)and a radial direction R. While the illustrated example is a high-bypassturbofan engine, the principles of the present invention are alsoapplicable to other types of engines, such as low-bypass turbofans,turbojets, turboprops, unducted fan engines or open rotor engines, etc.,as well as turbine engines having any number of compressor-turbinespools.

The engine 10 includes a fan 12, a low pressure compressor (“LPC”) orbooster 14, a high-pressure compressor or “HPC” 16, a combustion sectionor combustor 18, a high pressure turbine or “HPT” 20, and a low-pressureturbine or “LPT” 22, arranged in serial flow relationship. Collectively,the fan 12, LPC 14, and LPT 22 define a low-pressure system orlow-pressure spool of the engine 10. Collectively, the HPC 16 and HPT 20define a high-pressure spool of the engine 10. The high-pressure spooland combustor 18 may be referred to as a “core” or “core engine”.

In operation, pressurized air exiting the HPC 16 is mixed with fuel inthe combustor 18 and ignited, thereby generating combustion gases. Somework is extracted from these gases by the HPT 20 which drives the HPC 16via a high pressure shaft 24. The combustion gases then flow into theLPT 22, which drives the fan 12 and LPC 14 via a low pressure shaft 26.As used herein, the engine 10 is considered to be “operating” when fuelis being is supplied to and burned in the combustor, and the resultingcombustion gases are driving rotation of at least the core. As usedherein, the engine 10 is considered to be “shut down” when fuel is notbeing supplied to the combustor. It will be understood that “operating”encompasses numerous operating conditions having varying rotor speedsand varying thrust and/or power outputs. It will be understood that oneor more the rotors of the engine 10 may be rotating when fuel is notbeing provided. This may occur, for example because of wind passingthrough the engine 10 on the ground, relative wind passing through theengine during aircraft flight (i.e., “wind milling”), or rotation bytorque applied from a starter or similar apparatus. Rotation of theengine 10 using a starter (pneumatic, hydraulic, electric, or other) orusing an electric machine prior to igniting the engine 10, or prior toaccelerating the engine to a lightoff speed (to ignite the engine 10)may be referred to as “motoring” the engine 10.

The HPC 16 includes a number of stages of rotating blades and stationaryvanes, all surrounded by a compressor casing 28. The compressor casing28 incorporates a compressor bleed plenum 29 in fluid communication withthe compressor flowpath and in fluid communication with the exterior ofthe compressor casing 28 through at least one compressor bleed port 30.The compressor bleed plenum 29 may extend over all or a portion of thecircumference of the compressor casing 28. It will be understood thatdifferent engines may include one or more bleed ports and a particularengine may or may not include a bleed plenum of the type illustrated. Asused herein, the term “compressor bleed port” is used generically torefer to a port, opening, plenum, or passage in compressor casing 28 orother analogous structure, such as a combustor casing (e.g., locateddownstream of the compressor, that is directly or indirectly in fluidcommunication with the compressor flowpath). The term “compressor bleedport” may refer to an existing structure within the engine 10, or astructure that is newly added to accommodate the reverse bleed systemdescribed herein.

One or more bleed ducts 32 are coupled to the compressor bleed port 30and are configured to conduct extracted airflow away from the HPC 16.The extracted air may be vented for the purpose of controlling thecompressor operating line, or may be specifically added for the purposeof introducing reverse bleed cooling air, as described herein.Alternatively, it may be used for purposes such as environmental controlsystems (“ECS”), pneumatically-powered actuators, engine hot sectioncooling, and/or clearance control systems. The bleed duct 32 may includea bleed control valve 34 operable to move between open and closedpositions, thereby controlling flow through the compressor bleed port30.

The engine core is surrounded by (i.e., is contained within) a core cowl36 or core nacelle, which defines an inboard boundary of a bypassflowpath 38 over which fan bypass air flows. Shown is a ducted turbofan,which includes the fan 12 surrounded by a fan nacelle 37 which isspaced-away from the core cowl 36 and defines an outboard boundary ofthe bypass flowpath 38. In this example the bypass flowpath 38 couldalso be referred to as a “fan duct”.

It will be appreciated, however, that in other exemplary embodiments,the engine 10 may not include the fan nacelle 37, and instead may be an“open rotor” turbofan engine (or other type of engine). With such aconfiguration, the bypass flowpath 38 would be bounded only by an outersurface of the core cowl 36.

The space inboard of the core cowl 36 and outboard of a core airflowpath is generally referred to as “undercowl space” 40. In practice,the undercowl space 40 may be vented to ambient external environment forexample through a vent 41 (shown schematically in FIG. 1 ). Things thatare said to be internal to the engine, for purposes of this disclosure,means things that are located within the space surrounded by the fannacelle 37, or the core cowl 36 (in a case such as an open rotor enginewhere the fan nacelle 37 is not present).

The engine 10 may optionally incorporate a variable bleed valve (“VBV”)system for controlling LPC stall margin. The VBV system includes one ormore variable bleed valves 42 mounted within a fan hub frame 44. Thevariable bleed valves 42 may be open during low power operation of theengine 10, such as at idle, for bleeding a portion of the compressedair. The variable bleed valves 42 are closed at high power operation ofthe engine 10, such as during cruise or takeoff, since bleeding is nolonger required. When the variable bleed valves 42 are open, air ispassed from the LPC flowpath through the fan hub frame 44 and into,e.g., the bypass passage 38 or other bypass space external of thenacelle 37. In the illustrated example, the engine 10 includes at leastone bypass duct 46 defining an air flowpath from the fan hub frame 44 toa bleed vent 48 communicating with the bypass flowpath 38.

Referring now also to FIG. 2 , providing a close-up view of, e.g., theHPC 16 and combustor 18 of FIG. 1 , the combustor 18 includes aplurality of fuel nozzles 50 which are supplied during engine operationwith pressurized liquid fuel. The fuel nozzles 50 are connected to afuel system 52 operable to supply a flow of pressurized liquid fuel atvarying flowrates according to operational need. As depictedschematically, the fuel system 52 supplies fuel through a fuel valve 54coupled to a fuel conduit 56, which is in turn coupled the fuel nozzles50. In some embodiments, the fuel nozzles 50 and fuel system 52 mayimplement more than one independent fuel flow circuit (e.g., pilot andmain circuits).

It will be understood that each fuel nozzle 50 may generally be ametallic mass including numerous small passages and orifices. The fuelnozzles 50 are subject to the formation of carbon (or “coke”) depositswhen a hydrocarbon fuel is exposed to high temperatures in the presenceof oxygen. This process is referred to as “coking” and is depending on,e.g., an oxygen content of the fuel, coking may generally be a risk whentemperatures exceed about 170 degrees C. (350 degrees F.).

During engine operation, both fuel and compressed air flow through thefuel nozzles 50, and the fuel nozzles 50 are bathed in an external flowof relatively cool compressor discharge air. All of these flows carryaway heat from the fuel nozzles 50, keeping fuel temperatures relativelylow. More specifically, the relatively high volume of fuel through thefuel nozzles 50 primarily maintains a temperature of the fuel nozzles 50at a relatively low temperature.

When engine operation stops, a volume of fuel may remain in the fuelnozzles 50 and may be heated to coking temperatures. Small amounts ofcoke interfering with fuel flow through the orifices in the fuel nozzles50 can make a large difference in fuel nozzle performance.

As will also be appreciated, when engine operation stops, the flow ofcompressor discharge air also stops. For example, the various turbinesection components, which are consistently exposed to combustion gassesat relatively high temperatures, may remain relatively hot following ashutdown of the engine (when the engine operation stops). The heat fromthese relatively hot components may conduct along the high pressurespool and into the combustor 18. The heat may also move generallyupward. When the compressor discharge air is flowing, the airflow maymaintain a relatively constant temperature in a circumferentialdirection of the engine. Once this airflow stops, however, a thermalmismatch between a upper portion of the spool and a lower portion of thespool may form (as the compressor discharge air is no longer providingfor the constant circumferential temperature gradient), creating a “bow”in the spool, also referred to as a “rotor bow.”

The inventors' analysis and testing has shown that if a flow of air atan appropriate pressure and flow rate is provided to the compressorsection, the combustion section, or both (e.g., back through thecompressor bleed port 30) following a shutdown of the engine 10, thisflow (“a reverse bleed”) can preferentially flow downstream from the HPC16 and provide cooling to the fuel nozzles 50 so as to reduce or preventfuel nozzle coking. In addition, such a flow may provide cooling to thecomponents susceptible to rotor bow to reduce an amount of rotor bow inthe engine 10. For example, at least a portion of this reverse bleed mayflow through the compressor section to reduce a rotor bow.

In particular, for the exemplary embodiment depicted, the gas turbineengine further includes an air cooling system in selective airflowcommunication with the compressor section, the combustion section, orboth for providing a flow of cooling air over the fuel nozzles 50 duringcertain operations, such as during a shutdown of the engine 10 or aftera shutdown of the engine 10. For the embodiment shown, the air coolingsystem is in airflow communication with a bleed air assembly (and inparticular the compressor bleed port 30 for the embodiment shown), andas such may be referred to as a reverse bleed system 60.

As will be appreciated, however, in other embodiments, the air coolingsystem may be any other suitable air cooling system for generating aflow of cooling air into or through the compressor section, thecombustion section, or both to reduce coking and/or rotor bow.

More specifically, as is depicted in FIG. 2 , the reverse bleed system60 may be used to supply cooling air flow, or a reverse bleed flow overthe fuel nozzles 50 during a shutdown of the engine 10 or after ashutdown of the engine 10.

The reverse bleed system 60 includes a cooling duct 62 disposed in theengine 10. It may be mounted, for example, wholly or partially in theundercowl space 40. In particular, for the embodiment shown, it ismounted completely within the undercowl space 40, “internal to” theundercowl space 40. The cooling duct 62 defines an inlet 64 disposed influid communication with a source of cooling air and an outlet 66disposed in airflow communication with the compressor section (such asthe HPC 16), the combustor 18, or both. Particularly for the embodimentshown, the cooling duct 62 is in airflow communication with the HPC 16and combustor 18 via the compressor bleed port 30. The complete coolingduct 62 may be built up from components such as tubes, connectors, pipejoints, and the like.

In the embodiment of FIG. 2 , the inlet 64 is connected in fluidcommunication with the bypass duct 46. In the embodiment of FIG. 2 , theoutlet 66 is connected to the existing bleed duct 32, which in turn isconnected to the compressor bleed port 30.

However, in other embodiments, the cooling duct 62 may be configured inany other suitable manner to provide a reverse bleed flow over the fuelnozzles 50. For example, the cooling duct 62 may be directly connectedto a dedicated opening in the cowl 36, may receive environmental airwithin the undercowl space 40, etc. For example, as is depicted inphantom, the cooling duct 62 may not extend to the bypass duct 46, andinstead the cooling duct 62 may simply be open to the undercowl space40, thereby allowing air from the undercowl space 40 to be directlydrawn into the valve assembly 68 rather than from the bypass duct 46.With such a configuration, the inlet 64 may accordingly be exposed tothe undercowl space 40.

Referring still to FIG. 2 , the cooling duct 62 incorporates a valveassembly 68 including one or more valves operable to control airflowbetween the inlet 64 and the outlet 66. Two or more valves may be usedto provide redundancy, and/or monitor or control airflow through duct62. In this example, first and second valves 70, 72 are used in series,where the first valve 70 is closest to the outlet 66. Stated anotherway, the first and second valves 70, 72 are in “series fluidcommunication”, meaning that a fluid flow passes through one valvebefore encountering another valve. “Series flow communication” stands incontrast to “parallel flow communication”.

In the illustrated example, the first valve 70 is a check valve whichmay be passively biased towards an open position by, e.g., a spring,stored fluid pressure, weight, or other suitable mechanism and arrangedto permit airflow in a direction from the inlet 64 towards the outlet66, but to block airflow in the opposite direction. It will beunderstood that valves may exhibit some fluid leakage even in the closedposition. Accordingly, the operation of a valve in the closed positionto block airflow, except for inherent leakage, may be described as“substantially preventing flow.”

In the illustrated example, the second valve 72 is a controllable valvehaving a flow control element (e.g., a gate, flapper, ball, etc.)movable between open and closed positions. In the open position, thesecond valve 72 permits airflow between the inlet 64 and the outlet 66.In the closed position, the second valve 72 blocks airflow between theinlet 64 and the outlet 66.

Numerous types of controllable valves may be used. In one example, thecontrollable valve may incorporate or be coupled to an actuator 74 whichprovides motive force for the valve's flow control element. Examples ofsuitable types of actuators include pneumatic, hydraulic, or electricaldevices.

In one example, the controllable valve may be of a type in which aspring or similar element urges the controllable valve towards an openposition, and fluid pressure acts in opposition to the spring to movethe valve towards the closed position. Suitable fluids could include,for example compressed air, pressurized oil, or pressurized fuel. In oneexample, the controllable valve may be coupled to the fuel system 52described above (see FIG. 1 ) in such a manner that pressurized fuel maybe provided to the valve during engine operation. The fuel pressure thustends to keep the valve closed when the engine 10 is operating. Thistype of valve may be referred to as a fluid-pressure-responsive passivevalve, for example a “passive fuel valve”.

In this specific example where one of the first and second valves 70, 72is a check valve and the other of the first and second valves 70, 72 isa controllable valve, either valve may be placed in the upstream ordownstream position relative to the other valve. However, check valvestend to close more reliably when subjected to a greater pressuredifferential. The first valve 70 would inherently be exposed to a higherair pressure, being closer to the compressor bleed port 30. Accordingly,the first valve 70 may be a check valve.

The cooling duct 62 includes a cooling blower 76 between the valveassembly 68 and the inlet 64. The cooling blower 76 may be any apparatusoperable to blow, pump, or move a cooling airflow from the inlet 64towards the outlet 66. In the illustrated example, the cooling blower 76includes a rotor 78 carrying a plurality of fan blades. The blower 76may, in the alternative, be located at, within or proximate to the inlet64 and distal of the valves 72, 70.

A power source for operating the cooling blower 76 may be mechanical,hydraulic, pneumatic, or electrical. In the illustrated example, theblower's rotor 78 is coupled to an electric motor 80. In one example,the motor 80 may be an AC induction motor or DC motor.

The cooling blower 76 may be sized to provide an adequate dischargepressure and flow rate for the cooling process described in more detailbelow. As one example, the cooling blower 76 may be sized to produce airflow on the order of approximately 0.05 kg/s (0.1 lb/s) to approximately0.23 kg/s (0.5 lb/s) at approximately 0.69 kPa (0.1 psi) toapproximately 6.9 kPa (1 psi). In one exemplary end use, the coolingblower 76 may be sized to produce air flow on the order of approximately0.12 kg/s (0.25 lb/s) at approximately 3.4 kPa (0.5 psi).

Operation of the reverse bleed system 60 is generally as follows. Whenthe engine 10 is running, the reverse bleed system 60 is inactive.Portions of the cooling air duct 62 will be pressurized withhigh-temperature air coming from the compressor bleed port 30. The valveassembly 68 will block the majority of the flow from the outlet 66towards the inlet 64. As noted above, some valve leakage is expected tooccur. Any leakage will pass through the cooling blower 76, inlet 64,and in the example of FIG. 2 , through the bypass duct 46 and vent 48.

After engine shutdown, soakback may occur which may heat the fuelnozzles 50 to an unacceptable temperature, and further certaincomponents of the engine 10 may experience rotor bow. The reverse bleedsystem 60 may be used to move cooling air flow from the inlet 64 throughthe cooling duct 62, through the outlet 66 and into the compressor bleedport 30. Subsequently, the cooling air can pass over the fuel nozzles 50and other parts of the core to lower their respective temperatures andto reduce or prevent coking and to reduce or prevent rotor bow.Fundamentally, the reverse bleed system 60 is employed by (1) operatingthe cooling blower 76 and (2) opening the valve or valves of the valveassembly 68, at a time during or after a shutdown of the engine 10. Forexample, the reverse bleed system 60 may be operated following ashutdown of the engine 10 and prior to a subsequent startup of theengine 10.

As a possible alternative, the reverse bleed system 60 could be used tomove cooling air flow from downstream portions of the engine 10, throughthe compressor bleed port 30, through the outlet 66, through the coolingduct 62, and out through the inlet 64. In this sense the so-called“reverse bleed” system 60 would be used to cause air movement throughthe bleed port in the same direction as airflow through the bleed portduring flight. This could be accomplished by assuring that all valvesare open or otherwise configured to permit flow in this direction andoperating the cooling blower 76 to move air in the opposite direction asdescribed above. Stated another way, the cooling blower 76 could be usedto “suck” air from the engine 10 rather than “blow” it into the engine10.

It will be appreciated, however, that in other exemplary embodiments,the engine 10 may include any other suitable air cooling system forproviding a cooling airflow over the fuel nozzles 50, or otherwisecapable of cooling the fuel nozzles 50 and other components susceptibleto rotor bow. For example, the air cooling system may be configured toprovide cooling air from any suitable location (e.g., ambient,under-cowl 40, compressor section, dedicated cooling airflow source,etc.). Additionally, or alternatively, the air cooling system may beconfigured to utilize the cooling airflow to reduce a temperature of thefuel nozzles 50 or other component susceptible to coking from soak back(e.g., certain fuel lines), component causing bowed rotor, etc. in anyother suitable manner. For example, the air cooling system may beconfigured to provide the cooling airflow over the components directly,may be configured to cool components through an intermediate component(e.g., cooling components thermally coupled to the components to becooled), etc. Additionally, or alternatively, still, the air coolingsystem may be positioned at any other suitable location for performingthe functions described herein.

In addition, numerous arrangements are possible for control andoperation of the air cooling system/the reverse bleed system 60. Inparticular, back also to FIG. 1 , the exemplary gas turbine engine 10further includes an engine controller 82, such as a Full AuthorityDigital Engine Control controller (“FADEC”) or Electronic EngineController (“EEC”). The engine controller 82 is configured to receivethe data sensed from one or more sensors and, e.g., make controldecisions based on the received data. In the embodiment depicted, theengine 10 includes sensors for sensing data indicative of engine speeds,engine temperatures, etc. In particular, the exemplary engine 10 shownincludes a first sensor 84 for sensing data indicative of a rotationalspeed of the low speed spool, a second sensor 86 for sensing dataindicative of a rotational speed of the high speed spool, and a thirdsensor 88 for sensing data indicative of an engine temperature (and inparticular of a turbine inlet temperature), and a fourth sensor 90 forsensing data indicative of another engine temperature (and in particularof an exhaust gas temperature). It will be appreciated that each ofthese sensors may be a single sensor, or an array of sensors, may be anysuitable type of sensor for sensing the data indicative of theparameter, and further may be located at any suitable location forsensing the data indicative of the parameter.

Referring particularly to the operation of the controller 82, in atleast certain embodiments, the controller 82 can include one or morecomputing device(s) 92. The computing device(s) 92 can include one ormore processor(s) 92A and one or more memory device(s) 92B. The one ormore processor(s) 92A can include any suitable processing device, suchas a microprocessor, microcontroller, integrated circuit, logic device,and/or other suitable processing device. The one or more memorydevice(s) 92B can include one or more computer-readable media,including, but not limited to, non-transitory computer-readable media,RAM, ROM, hard drives, flash drives, and/or other memory devices.

The one or more memory device(s) 92B can store information accessible bythe one or more processor(s) 92A, including computer-readableinstructions 92C that can be executed by the one or more processor(s)92A. The instructions 92C can be any set of instructions that whenexecuted by the one or more processor(s) 92A, cause the one or moreprocessor(s) 92A to perform operations. In some embodiments, theinstructions 92C can be executed by the one or more processor(s) 92A tocause the one or more processor(s) 92A to perform operations, such asany of the operations and functions for which the controller 82 and/orthe computing device(s) 92 are configured, the operations for operatinga gas turbine engine 10 and/or cooling system/reverse bleed system 60(e.g., method 300), as described herein, and/or any other operations orfunctions of the one or more computing device(s) 92. The instructions92C can be software written in any suitable programming language or canbe implemented in hardware. Additionally, and/or alternatively, theinstructions 92C can be executed in logically and/or virtually separatethreads on processor(s) 92A. The memory device(s) 92B can further storedata 92D that can be accessed by the processor(s) 92A. For example, thedata 92D can include data indicative of power flows, data indicative ofengine/aircraft operating conditions, and/or any other data and/orinformation described herein.

The computing device(s) 92 can also include a network interface 92E usedto communicate, for example, with the other components of the gasturbine engine 10, the aircraft incorporating the gas turbine engine,etc. For example, in the embodiment depicted, as noted above, the gasturbine engine 10 includes one or more sensors 84, 86, 88, 90 forsensing data indicative of one or more parameters of the gas turbineengine. The controller 82 is operably coupled to the one or more sensorsthrough, e.g., the network interface 92E, such that the controller 82may receive data indicative of various operating parameters sensed bythe one or more sensors during operation. Further, for the embodimentshown the controller 82 is operably coupled to, e.g., the air coolingsystem/reverse bleed system 60. In such a manner, the controller 82 maybe configured to operate the reverse bleed system 60 in response to,e.g., the data sensed by the one or more sensors.

The network interface 92E can include any suitable components forinterfacing with one or more network(s), including for example,transmitters, receivers, ports, controllers, antennas, and/or othersuitable components.

The technology discussed herein makes reference to computer-basedsystems and actions taken by and information sent to and fromcomputer-based systems. One of ordinary skill in the art will recognizethat the inherent flexibility of computer-based systems allows for agreat variety of possible configurations, combinations, and divisions oftasks and functionality between and among components. For instance,processes discussed herein can be implemented using a single computingdevice or multiple computing devices working in combination. Databases,memory, instructions, and applications can be implemented on a singlesystem or distributed across multiple systems. Distributed componentscan operate sequentially or in parallel.

Referring now briefly to FIG. 3 , providing a perspective, schematicview of an aircraft 100 as may incorporate the exemplary engine 10 andreverse bleed system 60 described above, the engine 10 incorporatingreverse bleed system 60 may be installed in the aircraft 100 having atleast one electrical power source such as a battery and inverter 102, anauxiliary power unit 104, a connection to a ground power unit 106 orother ground-based power source, or another engine 10 having anelectrical generator.

In one example, the aircraft 100 includes an electronic aircraftcontroller 108 in data communication with the engine controller 82described above (see FIG. 1 ) and also including a controllableelectrical power connection 110 to the air cooling system/reverse bleedsystem 60 (including, e.g., the cooling blower 76). The aircraftcontroller 108 may have connections to various inputs such as cockpitswitch positions, an/or sensors such as an outside air temperature (OAT)probe 112 or a weight-on-wheels sensor 114. The aircraft controller 108may be configured in a similar manner as the exemplary engine controller82 described above with reference to FIG. 1 .

Various control methods in accordance with exemplary aspects of thepresent disclosure are described below with reference to FIG. 4 .

In particular, referring now to FIG. 4 , a method 200 for operating agas turbine engine is provided. In certain exemplary aspects, the method200 may be utilized with one or more of the exemplary aircraft, engines,air cooling systems (e.g., reverse bleed systems), etc. described abovewith reference to FIGS. 1 through 3 . However, in other exemplaryaspects, the exemplary method 200 may be utilized with any othersuitable aircraft, engine, air cooling systems, etc.

For the exemplary aspect of the method 200 depicted in FIG. 4 , themethod 200 generally includes at (202) determining data indicative of anoperation of an air cooling system of the gas turbine engine during ashutdown of the gas turbine engine, following the shutdown of the gasturbine engine, or both, and at (204) modifying a start-up schedule ofthe gas turbine engine in response to the determined data indicative ofthe operation of the air cooling system of the gas turbine engine.

The term “startup schedule” generally refers to the startup operatingprocedure for a gas turbine engine, including in certain exemplaryaspects, an amount of time to rotate the engine with a starter or otherelectric machine prior to increasing the rotational speed of the engineto a lightoff rotational speed.

In particular, for the exemplary aspect of the method 200 depicted inFIG. 4 , the air cooling system may be any suitable air cooling systemfor providing a flow of cooling air to a component of the gas turbineengine configured to contain fuel during or after a shutdown of theengine, one or more components susceptible to rotor bow, or both. Morespecifically, for the exemplary aspect of the method 200 depicted inFIG. 4 , the air cooling system is a reverse bleed system configured toprovide a flow of cooling air over a component of a combustor of the gasturbine engine during operation of the reverse bleed system. Forexample, the reverse bleed system may be configured to provide a flow ofcooling air over one or more fuel nozzles of the combustor of the gasturbine engine during operation of the reverse bleed system.

More specifically, it will be appreciated that prior to determining thedata indicative of the operation of the air cooling system/reverse bleedsystem at (202) and modifying the start-up schedule at (204), for theexemplary aspect of the method 200 depicted in FIG. 4 , the method 200additionally includes at (206) initiating operation of the reverse bleedsystem following the shutdown of the gas turbine engine. Initiating theoperation of the reverse bleed system at (206) may be at least in partin response to one or more environmental/ambient condition parametersand engine parameters, and further may be a command from an enginecontroller.

For example, for the exemplary aspect depicted, the method 200 furtherincludes at (208) determining data indicative of a soakback temperatureindicator parameter. The soakback temperature indicator parameter may beany suitable parameter for indicating that an anticipated temperature ofone or more components configured to contain fuel following a shutdownof the engine will exceed a predetermined threshold. The predeterminedthreshold may be a temperature threshold at which any fuel in thecomponent is likely to coke. For example, the soakback temperatureindicator parameter may be an ambient condition parameter, such as anambient temperature parameter, an ambient altitude parameter, etc.Additionally, or alternatively, the soakback temperature indicatorparameter may be an engine temperature parameter, such as an exhaust gastemperature, compressor exit temperature, turbine inlet temperature,etc. The soakback temperature indicator parameter may additionally oralternatively be based on software heat transfer models utilizing one ormore of the above parameters and/or other data, etc.

For the exemplary aspect depicted, initiating operation of the reversebleed system at (206) further includes at (210) initiating operation ofthe reverse bleed system following the shutdown of the gas turbineengine in response to the data indicative of the soakback temperatureindicator parameter determined at (208). For example, the datadetermined at (208) may include data indicative of the soakbacktemperature indicator parameter exceeding a predetermined threshold, andthe method 200 may initiate operation of the reverse bleed system inresponse to such data.

Also by way of example, for the exemplary aspect depicted the method 200further includes at (212) determining data indicative of an engineoperating parameter. The engine operating parameter may be any suitableparameter indicating that the engine and/or aircraft incorporating theengine is in a desired operating condition for initiating the reversebleed system. For example, the data indicative of the engine operatingparameter may be data indicative of the engine being powered on or off(e.g., from a user/operator switch within a cockpit of the aircraft, oran electronic signal), a rotational speed of one or more components ofthe engine, a weight on wheels sensor reading (e.g., from a weight onwheels sensor 114, to ensure the shutdown is not a mid-flight shutdown),etc. Additionally or alternatively, the data indicative of the engineoperating parameter may include data indicative of various otherconditions of systems of the engine, such as an open/closed indicator ona reverse bleed valve of the engine.

In one exemplary aspect of the method 200, as will be described in moredetail below with reference to, e.g., FIG. 5 , in certain exemplaryaspects, the data indicative of the engine operating parameter mayinclude data indicative of a rotational speed of a shaft of the engine,such as a high-pressure shaft of the engine. The data indicative of theengine operating parameter may include data indicative of a rate ofdecay of the rotational speed of the shaft of the engine and/or dataindicative of the rotational speed falling below a predeterminedthreshold (e.g., 10% of the engine's rated speed, such as 5% of theengine's rated speed).

With such an exemplary aspect, it will be appreciated that initiatingoperation of the reverse bleed system at (206) further includes at (214)initiating operation of the reverse bleed system following the shutdownof the gas turbine engine in response to the data indicative of theengine operating parameter determined at (212). For example, in certainexemplary aspects, e.g., where the data indicative of the engineoperating parameter may include data indicative of a rate of decay ofthe rotational speed of the shaft of the engine and data indicative ofthe rotational speed falling below a predetermined threshold, initiatingoperation of reverse bleed system at (214) may include initiatingoperation of the reverse bleed system after a determined amount of timefollowing the engine falling below the predetermined threshold. Thedetermined amount of time may be a preset time (e.g., a predeterminedamount of time), or alternatively may be based on the rate of decay ofthe rotational speed of the shaft.

Following initiating operation of the reverse bleed system at (206), themethod 200 further includes at (216) operating the reverse bleed systemfor an amount of time. Operating the reverse bleed system at (216) mayinclude operating the reverse bleed system for a determined amount oftime (e.g., based on one or more sensed parameters, based on thesoakback temperature indicator parameter (e.g., high ambienttemperature, longer operation; higher engine temperature, longeroperation; higher altitude, longer operation), etc.), or alternatively,may include operating the reverse bleed system for a predeterminedamount of time (e.g., 30 minutes, 60 minutes, 90 minutes, etc.).Operating the reverse bleed system at (216) includes at (217) providinga flow of cooling air through a high-pressure compressor of the gasturbine engine, the combustion section gas turbine engine, or both.

For example, in certain aspects, providing the flow of cooling air at(217) may include providing the flow of cooling air through thehigh-pressure compressor, the combustion section, or both and over oneor more components of the gas turbine engine configured to contain fuelfollowing the shutdown of the gas turbine engine.

Alternatively, however, providing the flow of cooling air at (217) mayinclude extracting air from the high-pressure compressor, the combustionsection, or both to generate a flow of cooling air over one or morecomponents of the gas turbine engine configured to contain fuelfollowing the shutdown of the gas turbine engine. For example, the oneor more components may be one or more fuel nozzles of the combustionsection of the gas turbine engine.

Notably, in certain exemplary aspects of the method 200, the method 200may further include terminating operation of the reverse bleed systemprior to the determined amount of time or prior to the predeterminedamount of time. The operation of the reverse bleed system may beterminated prior to the determined or predetermined amount of time as aresult of a restarting of the gas turbine engine, disconnecting the gasturbine engine from a power source (such as a ground power source), acommanded termination for maintenance operations, etc. With suchexemplary aspect, the method 200 may save data indicative of the amountof time the reverse bleed system operated, which may be indicative ofthe operation of the reverse bleed system (and used, e.g., at step(202)).

It will be appreciated that in at least certain exemplary aspects, anengine controller for the engine including the reverse bleed system maybe configured to power down after an amount of time following theshutdown of the engine. In certain exemplary aspects, this amount oftime may be less than the amount of time it is desirable to operate thereverse bleed system. In such a manner, it will be appreciated that forthe exemplary aspect of the method 200 depicted, operating the reversebleed system for the amount of time at (216) further includes at (218)powering the reverse bleed system with a power source external to thegas turbine engine. For example, in certain exemplary aspects, the powersource external to the gas turbine engine may be an electric energystorage unit of the aircraft (e.g., a battery pack), a ground powersystem, an auxiliary power unit of the aircraft, an electric machinecoupled to or driven by another engine of the aircraft, etc.

More specifically, for the exemplary aspect of the method 200 depictedin FIG. 4 , it will be appreciated that the method 200 further includesat (220) shutting down the engine controller of the gas turbine engineafter a first amount of time following the shutdown of the gas turbineengine and following initiating operation of the reverse bleed system at(206). With such an exemplary aspect, operating the reverse bleed systemfor the amount of time at (216) further includes at (222) operating thereverse bleed system for a second amount of time following the shutdownof the gas turbine engine, where the second amount of time is greaterthan the first amount of time. For example, the second amount of timemay be at least 50% greater, such as 100% greater, such as five timesgreater, such as up to 100 times greater than the first amount of time.

As briefly noted above, the method 200 further includes at (202)determining data indicative of the operation of the gas turbine engineduring the shutdown of the gas turbine engine, following the shutdown ofthe gas turbine engine, or both. In certain exemplary aspects, the dataindicative of the operation of the reverse bleed system determined at(202) may include data indicative of the reverse bleed system operatingproperly.

For example, in certain exemplary aspects, the data indicative of theoperation of the reverse bleed system determined at (202) may includedata indicative of an engine temperature at a first time following theinitiation of the reverse bleed system at (206) and data indicative ofthe engine temperature at a second time following initiation of thereverse bleed system at (206). The second time may be after the firsttime. Further, the first time may be a relatively short time periodafter initiating ration of the reverse bleed system at (206) and thesecond time may be relatively short time period prior to shutting downthe engine controller at (220). The “short time period” may refer toamount time less than or equal to about three minutes, such as less thanor equal to about one minute, such as less than or equal to about 30seconds, such as less than or equal to about 10 seconds.

With such an exemplary aspect, the data indicative of the operation ofthe reverse bleed system determined at (202) may further include dataindicative of a difference between the engine temperature at the firsttime and at the second time. For example, in determining the dataindicative of the operation of the reverse bleed system at (202), themethod may determine a slope between the engine temperature at the firsttime and at the second time to determine if the engine temperature isincreasing or decreasing. If the engine temperature is increasing, suchmay indicate that the reverse bleed system is not operating properly,whereas if the engine temperature is decreasing, such may indicate thatthe reverse bleed system is operating properly. The engine temperaturemay be, e.g., an exhaust gas temperature, a turbine inlet temperature, acompressor exit temperature, etc.

As also noted above, the method 200 further includes at (204) modifyingthe startup schedule of the gas turbine engine in response to the dataindicative of the operation of the reverse bleed system of the gasturbine engine determined at (202). For the exemplary aspect depicted,modifying the startup schedule of the gas turbine engine at (204)further includes at (224) reducing a motoring time of the gas turbineengine.

As will be appreciated, when an engine experiences a rotor bow, it maybe necessary to “motor” the engine for an amount time prior toinitiating the remaining starting sequence for the gas turbine engine.In such manner, it will be appreciated that motoring generally refers torotating one more components of the engine with, e.g., a starter orother electric motor to allow the bowed components to distribute theheat to reduce the bowing of the component. The motoring process can betime consuming depending on the degree of the rotor bow. However, if itis determined that the reverse bleed system has been operating properly,such may indicate that the components are not “bowed” or are not bowedto the same extent that they otherwise would be, which may allow for areduction in the motoring time of the engine prior to initiating theremaining portion of the starting sequence for the gas turbine engine.

It will be appreciated that in other exemplary aspects, the method 200may additionally or alternatively determine data indicative of thereverse bleed system not operating properly, or not operating at 100%effectiveness at (202). With such an exemplary aspect, modifying thestartup schedule of the gas turbine engine at (204) may additionally oralternatively include increasing a motoring time of the gas turbineengine within the startup sequence.

It will be appreciated, however, that the exemplary aspect of the method200 described above with reference to FIG. 4 is by way of example only.In other exemplary aspects, any other suitable air cooling system may beutilized in place of the reverse bleed system discussed, and furtherthat in other exemplary aspects, the air cooling system may beconfigured to provide a flow of cooling air to additional or alternativecomponents of the gas turbine engine. For example, in other exemplaryaspects, the air cooling system may be, e.g., a system for providing aforced airflow through a main airflow path of the engine through aninlet to the engine (e.g., a standalone fan position at a forward end ofthe gas turbine engine, the aft end of the gas turbine engine, or both).Also by way of example, in other exemplary aspects, the component cooledby the air cooling system may be any other suitable component configuredto contain fuel following a shutdown of the gas turbine engine, e.g.,one or more fuel lines, auxiliary burners or combustors, etc.

Referring now to FIG. 5 , a graph 300 is provided depicting a rotationalspeed of a gas turbine engine and an engine temperature of the gasturbine engine over a time period from a shutdown of the gas turbineengine to a subsequent startup of the gas turbine engine. One exemplaryoperation of the method 300 is described below with reference to thegraph 300 of FIG. 5 .

As will be appreciated, the graph 300 generally includes a Y-axisrepresenting rotational speed (on the left side of the graph 300, Y-axis302) and engine temperature (on the right side of the graph 300, Y-axis304), and an X-axis 306 depicting time. A first line 308 is depictedrepresenting the rotational speed of the engine over the time period anda second line 310 is depicted representing the engine temperature overthe same time period. The rotational speed may be a shaft speed of thegas turbine engine, such as a high-pressure shaft speed or low-pressureshaft speed. The engine temperature may be an exhaust gas temperature, aturbine inlet temperature, a compressor exit temperature, etc.

At T0, the gas turbine engine is shutdown. Shutting down the engine mayinclude, e.g., operating a switch or other control mechanism within acockpit of the engine by an operator of the gas turbine engine/aircraftincluding the gas turbine engine. Further, shutting down the engine mayinclude shutting down a fuel flow to a combustor of the gas turbineengine. Following the shutdown of the gas turbine engine, the rotationalspeed of the gas turbine engine drops. The engine temperature similardrops initially, but may begin to climb due to the amount of heat storedwithin the various components of the gas turbine engine, given arelatively high thermal mass of these components, and given that theengine speed is slowing down and airflow through the gas turbine engineis correspondingly decreasing.

At or around shutdown/T0, an aircraft controller, an engine controller,or both may provide a command to an air cooling system, whichspecifically for the embodiment shown may be a reverse bleed system, toactivate the air cooling system/reverse bleed system. In particular, forthe aspects shown, the command to activate the reverse bleed system mayinitially be a failsafe launch command from the engine controller or theaircraft controller to initiate operation of the reverse bleed systemafter a predetermined amount of time if the reverse bleed system is notalready operating.

More specifically, for the graph 300 shown in FIG. 5 , the enginecontroller may further determine when the rotational speed of the gasturbine engine has reached a predetermined level, shown at T1, which maybe, e.g., 10% or less of a rated speed for the gas turbine engine, suchas 5% or less of the rated speed of the gas turbine engine. Thispredetermined level is equal to or higher than a level at which thesensors or other mechanisms that determine the rotational speedtypically cut out.

After determining the engine has reached the predetermined level, theengine controller may wait an amount of time prior to initiatingoperation of the reverse bleed system at T2. In particular, the enginecontroller may send a command to the aircraft controller to providepower to the cooling system for, e.g., a determined amount of time or apredetermined amount of time. In response, the aircraft controller maybe programmed to provide electrical power to the reverse bleed systemfor the specified amount of time and then to shut off electrical power.

The amount of time between T1 and T2 may be based on a rate of decay ofthe rotational speed of the gas turbine engine, and/or one or more knownconfigurations of the gas turbine engine. For example, if the gasturbine engine includes, e.g., hydraulic pumps or other accessorysystems coupled to the high-pressure spool, these components mayincrease an amount of drag on the high-pressure spool (resulting in aquicker decrease in the rotational speed). In such a case, the amount oftime between T1 and T2 may be reduced. The amount of time between T1 andT2 may be, e.g., less than or equal to five minutes, such as less thanor equal to three minutes, such as less than or equal to two minutes,such as less than one minute, such as less than or equal to 30 seconds,such as greater than or equal to five seconds.

During typical operations, the reverse bleed system may operate for anamount of time after initiating operation at T2. In particular, for theembodiment shown, the reverse bleed system is configured to operate fromT2 to T6. Notably, the engine controller is typically configured to shutdown an amount of time after shutdown of the gas turbine engine at T0that is less than the amount of time from T0 to T6. In the graph 300depicted in FIG. 5 , the engine controller is configured to shut down atT5. As mentioned above, it will therefore be appreciated that operatingthe reverse bleed system may include providing power from a power sourceexternal to the gas turbine engine.

In order to determine if the reverse bleed system is operatingproperly/has operated properly, the engine controller is configured todetermine the engine temperature at a first time shortly afterinitiating operation of the reverse bleed system, and is furtherconfigured to determine the engine temperature at a second time shortlyprior to shutting down the engine controller. The first time shortlyafter initiating operation of the reverse bleed system shown at T3 andthe second time shortly prior to the shutting down of the enginecontroller is shown at T4. The terms “shortly after” and “shortly prior”are simply terms used for convenience and do not require any inherentlimitations. In certain exemplary aspects, these time periods may be,e.g., between two seconds and 30 seconds.

As will be appreciated from the engine temperature line 310 in the graph300 of FIG. 5 , if the reverse bleed system is operating properly, aslope between the engine temperature at the first time T3 and the secondtime T4 is a negative slope, indicating that the engine temperature isdecreasing. Depicted in phantom is an alternative engine temperatureline 310′ starting at T2 illustrating the engine temperature from T2 onif the reverse bleed system does not operate properly. As is shown, ifthe reverse bleed system is not operating properly, a slope between theengine temperature at the first time T3 and the second time T4 is apositive slope indicating that the engine temperature is increasing.

Accordingly, once it comes time to start the gas turbine engine back up,depicted at T7, if the reverse bleed system operated properly, theengine temperature will be relatively low. By contrast, if the reversebleed system did not operate properly, the engine temperature will berelatively high. If the engine temperature is relatively high, such mayindicate that coking has occurred within, e.g., the fuel nozzles, andfurther may indicate that the engine is experiencing a relatively highdegree of rotor bow. If the engine is experiencing a relatively highdegree of rotor bow, it will be necessary to motor the gas turbineengine for a relatively long amount of time to allow the heat toredistribute and the relatively high degree of rotor bow to mitigate. Bycontrast, if the engine temperature is relatively low, and the engine isexperiencing a relatively low amount of rotor bow, then it may not benecessary to motor the engine for very long prior to accelerating theengine for startup.

For example, if the reverse bleed system operated properly following theprior shut down, such as in the embodiment depicted in the graph 300 ofFIG. 5 , then a time period for motoring the gas turbine engine, shownbetween T7 and T8 may be relatively low, prior to accelerating theengine for startup at T8. Such may result in a relatively low amount oftime between initiating the startup sequence at T7 and achieving alight-off rotational speed at T9. By contrast, if the reverse bleedsystem did not operate properly following the prior shut down, such asdepicted in phantom in the graph 300 of FIG. 5 , then a time period formotoring the gas turbine engine, shown between T7 and T10 (via phantomengine speed line 308′), prior to accelerating the engine for startup atT10 is relatively high. Such may result in a relatively high amount oftime between initiating the startup sequence at T7 and achieving alight-off rotational speed at T11.

In such manner, the engine controller, the aircraft controller, or bothmay determine data indicative of an operation of the rotor bow systemfollowing the prior shut down of the gas turbine engine, and in responsemay modify the startup sequence for the subsequent startup of the gasturbine engine. For example, in the event the data indicative of theoperation of the rotor bow system indicates that the rotor bow systemoperated properly, modifying the startup sequence for the subsequentstartup of the gas turbine engine may include reducing a motoring timeof the gas turbine engine, saving time and energy for the subsequentstartup sequence.

Moreover, it will be appreciated that in addition to the exemplary stepsoutlined above, an engine controller for an engine incorporating an aircooling system/reverse bleed system in accordance with one or moreexemplary aspects of the present disclosure may further be configured toperform additional functions. For example, referring now to FIG. 6 , aschematic diagram of a control system 400 in accordance with the presentdisclosure is provided. The control scheme 400 generally includes acontroller 402, which may be an engine controller. The controller 402 isoperable with the air cooling system/reverse bleed system forcontrolling operation of the system and receiving data from the systemengine and engine including the system.

The control scheme 400 further includes a configuration block 404,whereby the controller 402 may confirm the air cooling system/reversebleed system is installed. The configuration block 404 may check that aspecific wiring harness is connected and that the correct software isinstalled.

The control scheme 400 further includes at block 406 an output dataprocessing block, whereby the controller 402 may communicate data with,e.g., the aircraft controller, and at block 408 a communicationprotocols block, providing communication protocols between the enginecontroller and, e.g., the aircraft controller.

Further, the control scheme 400 includes at block 410 a menu mode block,allowing for manual operation of the air cooling system/reverse bleedsystem in response to, e.g., one or more user inputs. The menu modeblock may allow for maintenance operations of the air coolingsystem/reverse bleed system, operability checks of the air coolingsystem/reverse bleed system, etc.

Further, the control scheme 400 includes at block 412 an air coolingsystem/reverse bleed system fault monitoring and processing block. Theblock 412 may allow for the storing and processing of additionalinformation that may indicate non-operation of the air coolingsystem/reverse bleed system. The block 412 may incorporate somepersistence, such that multiple failures must be indicated before repairor replacement of the air cooling system/reverse bleed system isrequested.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

Further aspects are provided by the subject matter of the followingclauses:

A method for operating a gas turbine engine including: determining dataindicative of an operation of a cooling system of the gas turbine engineduring a shutdown of the gas turbine engine, following the shutdown ofthe gas turbine engine, or both; and modifying a startup schedule of thegas turbine engine in response to the determined data indicative of theoperation of the cooling system of the gas turbine engine.

The method of one or more of these clauses, wherein the cooling systemprovides a flow of cooling air to a component of the gas turbine engineconfigured to contain fuel.

The method of one or more of these clauses, wherein the cooling systemis a reverse bleed system configured to provide a flow of air over acomponent of a combustor of the gas turbine engine during operation ofthe reverse bleed system.

The method of one or more of these clauses, wherein the cooling systemcomprises a blower positioned within an undercowl location of the gasturbine engine.

The method of one or more of these clauses, wherein the blower isconfigured to provide the flow of air from the undercowl location of thegas turbine engine, or from a bypass valve.

The method of one or more of these clauses, wherein the cooling systemdefines an inlet exposed to the undercowl location of the gas turbineengine.

The method of one or more of these clauses, wherein the cooling systemis configured to provide a flow of air through a compressor bleed portusing a blower located internal to a cowl of the gas turbine engine.

The method of one or more of these clauses, wherein modifying thestartup schedule of the gas turbine engine comprises reducing a motoringtime of the gas turbine engine.

The method of one or more of these clauses, wherein the data determinedindicative of the operation of the cooling system comprises dataindicative of the cooling system operating properly.

The method of one or more of these clauses, further comprising:initiating operation of the cooling system following the shutdown of thegas turbine engine, and wherein the data determined indicative of theoperation of the cooling system comprises data indicative of an enginetemperature at a first time after operation of the cooling system hasbeen initiated and data indicative of the engine temperature at a secondtime, wherein the second time is after the first time.

The method of one or more of these clauses, wherein the data determinedindicative of the operation of the cooling system further comprises dataindicative of a difference between the engine temperature at the firsttime and at the second time.

The method of one or more of these clauses, further comprising:determining data indicative of a soakback temperature indicatorparameter; and initiating operation of the cooling system following theshutdown of the gas turbine engine in response to the data indicative ofthe soakback temperature indicator parameter.

The method of one or more of these clauses, further comprising:determining data indicative of an engine operating parameter; andinitiating operation of the cooling system following the shutdown of thegas turbine engine in response to the data indicative of the engineoperating parameter.

The method of one or more of these clauses, further comprising:initiating operation of the cooling system following the shutdown of thegas turbine engine; and operating the cooling system for a predeterminedamount of time.

The method of one or more of these clauses, wherein operating thecooling system for a predetermined amount of time comprises powering thecooling system with a power source external to the gas turbine engine.

The method of one or more of these clauses, further comprising:initiating operation of the cooling system following the shutdown of thegas turbine engine; shutting down an engine controller of the gasturbine engine after a first amount of time following the shutdown ofthe gas turbine engine; and operating the cooling system for a secondamount of time following the shutdown of the gas turbine engine, whereinthe second amount of time is greater than the first amount of time.

The method of one or more of these clauses, further comprising:initiating operation of the cooling system following the shutdown of thegas turbine engine; and operating the cooling system, wherein operatingthe cooling system comprises providing a flow of cooling air through ahigh pressure compressor of the gas turbine engine, a combustion sectionof the gas turbine engine, or both.

A method for operating a gas turbine engine comprising: receiving dataindicative of a soakback temperature indicator parameter; initiatingoperation of a cooling system of the gas turbine engine during ashutdown of the gas turbine engine, following the shutdown of the gasturbine engine, or both in response to the received data indicative ofthe soakback temperature indicator parameter; and operating the coolingsystem to provide a flow of cooling air through a high pressurecompressor of the gas turbine engine, a combustion section of the gasturbine engine, or both.

The method of one or more of these clauses, wherein the data indicativeof the soakback temperature indicator parameter comprises dataindicative of an ambient condition, data indicative of an enginetemperature parameter, or both.

The method of one or more of these clauses, wherein operating thecooling system comprises operating the cooling system for an amount oftime determined based at least in part on the soakback temperatureindicator parameter.

An aeronautical system comprising: a gas turbine engine comprising acompressor section, a combustion section and a turbine section arrangedin serial flow order, the gas turbine engine further comprising acooling system in selective airflow communication with the compressorsection, the combustion section, or both; and a control system inoperable communication with the cooling system, the control systemconfigured to: determine data indicative of an operation of the coolingsystem of the gas turbine engine during a shutdown of the gas turbineengine, following the shutdown of the gas turbine engine, or both; andmodify a startup schedule of the gas turbine engine in response to thedetermined data indicative of the operation of the cooling system of thegas turbine engine.

The aeronautical system of one or more of these clauses, wherein thecooling system provides a flow of cooling air to a component of the gasturbine engine configured to contain fuel.

The aeronautical system of one or more of these clauses, wherein thecooling system is a reverse bleed system configured to provide a flow ofair over a component of a combustor of the gas turbine engine duringoperation of the reverse bleed system.

The aeronautical system of one or more of these clauses, whereinmodifying the startup schedule of the gas turbine engine comprisesreducing a motoring time of the gas turbine engine.

The aeronautical system of one or more of these clauses, wherein thedata determined indicative of the operation of the cooling systemcomprises data indicative of the cooling system operating properly.

The aeronautical system of one or more of these clauses, wherein thecontroller is further configured to initiate operation of the coolingsystem following the shutdown of the gas turbine engine, and wherein thedata determined indicative of the operation of the cooling systemcomprises data indicative of an engine temperature at a first time afteroperation of the cooling system has been initiated and data indicativeof the engine temperature at a second time, wherein the second time isafter the first time.

A controller for a gas turbine engine, the controller comprising one ormore processors and memory, the memory storing instructions that whenexecuted by the one or more processors cause the gas turbine engine to:determine data indicative of an operation of a cooling system of the gasturbine engine during a shutdown of the gas turbine engine, followingthe shutdown of the gas turbine engine, or both; and modify a startupschedule of the gas turbine engine in response to the determined dataindicative of the operation of the cooling system of the gas turbineengine.

We claim:
 1. A gas turbine engine, comprising: a core cowl, a corecontained within the core cowl, including a compressor in fluidcommunication with a downstream combustor and a downstream turbine, thecompressor including a compressor bleed port, wherein an undercowl spaceis defined between the core cowl and the core; a cooling duct disposedat least partially in the undercowl space and having an inlet and anoutlet, wherein the cooling duct is in fluid communication with a sourceof cooling air and is further in fluid communication with the compressorbleed port; a valve assembly including at least one valve disposed inthe cooling duct; and a cooling blower disposed within the engine andoperable to move an air flow from the inlet of the cooling duct towardsthe outlet of the cooling duct and through the compressor bleed port. 2.The engine of claim 1, wherein the inlet of the cooling duct is in fluidcommunication with the source of cooling air, and wherein the outlet ofthe cooling duct is in fluid communication with the compressor bleedport.
 3. The engine of claim 2, wherein the inlet of the cooling duct isin fluid communication with the undercowl space.
 4. The engine of claim2, wherein a bypass duct is disposed upstream of the compressor, and theinlet of the cooling duct is in fluid communication with the bypassduct.
 5. The engine of claim 1, further including a bleed duct coupledto the compressor bleed port, wherein the outlet of the cooling duct iscoupled to the bleed duct.
 6. The engine of claim 1, wherein the valveassembly includes a first valve and a second valve in series fluidcommunication, the first valve being closer to the outlet of the coolingduct.
 7. The engine of claim 6, wherein at least one of the valves is acheck valve.
 8. The engine of claim 6, wherein at least one of thevalves is a fluid-pressure-responsive passive valve.
 9. The engine ofclaim 6, wherein the first valve is a check valve and the second valveis a controllable valve.
 10. The engine of claim 1, wherein the coolingblower is at least partially disposed within the cooling duct.
 11. Theengine of claim 1, wherein the cooling blower is electrically powered.12. A cooling duct assembly for a gas turbine engine, the gas turbineengine comprising a core cowl and a core contained within the core cowl,the core including a compressor having a compressor bleed port, the gasturbine engine defining an undercowl space between the core cowl and thecore, the cooling duct assembly comprising: a cooling duct configured tobe disposed at least partially in the undercowl space and having aninlet and an outlet, wherein the cooling duct is configured to be influid communication with a source of cooling air and the compressorbleed port when installed in the gas turbine engine; a valve assemblyincluding at least one valve disposed in the cooling duct; and a coolingblower disposed operable to move an air flow through the cooling duct toprovide an airflow through the compressor bleed port when installed inthe gas turbine engine and during operation.
 13. The cooling ductassembly of claim 12, wherein the inlet of the cooling duct isconfigured to be in fluid communication with the source of cooling air,and wherein the outlet of the cooling duct is configured to be in fluidcommunication with the compressor bleed port.
 14. The cooling ductassembly of claim 13, wherein the inlet of the cooling duct isconfigured to be in fluid communication with the undercowl space. 15.The cooling duct assembly of claim 12, wherein the gas turbine enginefurther includes a bleed duct coupled to the compressor bleed port,wherein the outlet of the cooling duct is configured to be coupled tothe bleed duct.
 16. The cooling duct assembly of claim 12, wherein thevalve assembly includes a first valve and a second valve in series fluidcommunication, the first valve being closer to the outlet of the coolingduct.
 17. The cooling duct assembly of claim 16, wherein at least one ofthe valves is a check valve.
 18. The cooling duct assembly of claim 16,wherein at least one of the valves is a fluid-pressure-responsivepassive valve.
 19. The cooling duct assembly of claim 16, wherein thefirst valve is a check valve and the second valve is a controllablevalve.
 20. The cooling duct assembly of claim 12, wherein the coolingblower is at least partially disposed within the cooling duct.